GOE 611 AIRFOIL (goe611-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
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Airfoil: GOE 611 AIRFOIL (goe611-il) Reynolds number: 1,000,000 Max Cl/Cd: 125.62 at α=4.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe611-il-1000000.txt Download as CSV file: xf-goe611-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 611 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 1.000 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.000 -0.1696 0.09936 0.09767 -0.0684 0.9368 0.0122
-9.750 -0.1670 0.09602 0.09421 -0.0693 0.9046 0.0122
-9.500 -0.1638 0.09257 0.09066 -0.0690 0.8818 0.0124
-9.250 -0.1567 0.09009 0.08810 -0.0695 0.8647 0.0125
-9.000 -0.1498 0.08748 0.08542 -0.0705 0.8518 0.0126
-8.750 -0.1430 0.08475 0.08263 -0.0719 0.8411 0.0128
-8.500 -0.1364 0.08185 0.07970 -0.0735 0.8324 0.0130
-8.250 -0.1303 0.07886 0.07666 -0.0754 0.8243 0.0133
-8.000 -0.1246 0.07563 0.07343 -0.0778 0.8169 0.0137
-7.750 -0.1213 0.07231 0.07006 -0.0807 0.8097 0.0141
-7.250 -0.0947 0.05988 0.05751 -0.0983 0.7970 0.0153
-7.000 -0.0796 0.05496 0.05248 -0.1025 0.7917 0.0153
-6.750 -0.0639 0.04992 0.04733 -0.1058 0.7862 0.0153
-6.500 -0.0538 0.04553 0.04283 -0.1077 0.7808 0.0157
-6.250 -0.0348 0.04362 0.04086 -0.1086 0.7755 0.0159
-6.000 -0.0144 0.04141 0.03856 -0.1097 0.7699 0.0164
-5.750 0.0071 0.03875 0.03576 -0.1107 0.7644 0.0172
-5.500 0.0347 0.03309 0.02972 -0.1115 0.7595 0.0189
-5.250 0.0558 0.02897 0.02530 -0.1116 0.7539 0.0190
-5.000 0.0750 0.02499 0.02103 -0.1116 0.7484 0.0187
-4.750 0.0989 0.02419 0.02019 -0.1121 0.7422 0.0200
-4.500 0.1244 0.02336 0.01926 -0.1123 0.7355 0.0210
-4.250 0.1539 0.02159 0.01705 -0.1114 0.7294 0.0233
-4.000 0.1761 0.01788 0.01309 -0.1114 0.7226 0.0219
-3.750 0.2030 0.01685 0.01185 -0.1113 0.7154 0.0247
-3.500 0.2302 0.01652 0.01147 -0.1116 0.7063 0.0261
-3.250 0.2602 0.01723 0.01195 -0.1112 0.6959 0.0288
-2.250 0.3690 0.01165 0.00556 -0.1108 0.6521 0.0298
-2.000 0.3964 0.01105 0.00488 -0.1107 0.6403 0.0304
-1.750 0.4239 0.01083 0.00463 -0.1108 0.6278 0.0318
-1.500 0.4515 0.01038 0.00410 -0.1107 0.6152 0.0318
-1.250 0.4788 0.01004 0.00367 -0.1106 0.6012 0.0318
-1.000 0.5062 0.00978 0.00334 -0.1106 0.5878 0.0321
-0.750 0.5337 0.00960 0.00311 -0.1106 0.5760 0.0328
-0.500 0.5611 0.00947 0.00292 -0.1106 0.5649 0.0337
-0.250 0.5887 0.00935 0.00273 -0.1106 0.5542 0.0342
0.000 0.6165 0.00925 0.00259 -0.1106 0.5440 0.0347
0.250 0.6441 0.00920 0.00248 -0.1107 0.5346 0.0352
0.500 0.6719 0.00917 0.00241 -0.1108 0.5252 0.0356
0.750 0.6998 0.00914 0.00233 -0.1109 0.5165 0.0360
1.000 0.7272 0.00909 0.00221 -0.1109 0.5075 0.0382
1.250 0.7553 0.00907 0.00217 -0.1111 0.4993 0.0400
1.500 0.7827 0.00912 0.00218 -0.1111 0.4910 0.0421
1.750 0.8106 0.00916 0.00218 -0.1112 0.4827 0.0444
2.000 0.8348 0.00782 0.00240 -0.1115 0.4749 0.7195
2.250 0.8621 0.00734 0.00252 -0.1112 0.4657 1.0000
2.500 0.8890 0.00747 0.00258 -0.1111 0.4571 1.0000
2.750 0.9154 0.00763 0.00266 -0.1110 0.4469 1.0000
3.000 0.9425 0.00775 0.00274 -0.1110 0.4381 1.0000
3.250 0.9691 0.00790 0.00284 -0.1109 0.4296 1.0000
3.500 0.9957 0.00805 0.00293 -0.1109 0.4201 1.0000
3.750 1.0223 0.00821 0.00304 -0.1108 0.4106 1.0000
4.000 1.0483 0.00839 0.00317 -0.1107 0.4002 1.0000
4.250 1.0744 0.00858 0.00331 -0.1106 0.3894 1.0000
4.500 1.1004 0.00876 0.00344 -0.1104 0.3788 1.0000
4.750 1.1259 0.00897 0.00360 -0.1102 0.3678 1.0000
5.000 1.1510 0.00921 0.00378 -0.1100 0.3558 1.0000
5.250 1.1757 0.00946 0.00397 -0.1096 0.3426 1.0000
5.500 1.1993 0.00978 0.00420 -0.1091 0.3233 1.0000
5.750 1.2213 0.01019 0.00448 -0.1083 0.3008 1.0000
6.000 1.2416 0.01070 0.00483 -0.1073 0.2719 1.0000
6.250 1.2597 0.01133 0.00525 -0.1058 0.2377 1.0000
6.500 1.2750 0.01211 0.00579 -0.1039 0.1986 1.0000
6.750 1.2817 0.01333 0.00665 -0.1005 0.1383 1.0000
7.000 1.2998 0.01382 0.00708 -0.0991 0.1280 1.0000
7.250 1.3180 0.01425 0.00748 -0.0976 0.1216 1.0000
7.500 1.3369 0.01457 0.00781 -0.0963 0.1188 1.0000
7.750 1.3554 0.01493 0.00819 -0.0949 0.1164 1.0000
8.000 1.3719 0.01539 0.00864 -0.0932 0.1122 1.0000
8.250 1.3880 0.01590 0.00914 -0.0915 0.1072 1.0000
8.500 1.4082 0.01621 0.00949 -0.0905 0.1046 1.0000
8.750 1.4245 0.01673 0.00999 -0.0890 0.0982 1.0000
9.000 1.4400 0.01731 0.01052 -0.0874 0.0863 1.0000
9.500 1.4287 0.02108 0.01390 -0.0786 0.0152 1.0000
9.750 1.4420 0.02193 0.01481 -0.0771 0.0141 1.0000
10.000 1.4541 0.02291 0.01585 -0.0755 0.0132 1.0000
10.250 1.4639 0.02407 0.01708 -0.0738 0.0123 1.0000
10.500 1.4704 0.02553 0.01864 -0.0719 0.0114 1.0000
10.750 1.4818 0.02668 0.01984 -0.0706 0.0111 1.0000
11.000 1.4919 0.02796 0.02118 -0.0693 0.0107 1.0000
11.250 1.5007 0.02937 0.02265 -0.0680 0.0102 1.0000
11.500 1.5082 0.03092 0.02428 -0.0667 0.0097 1.0000
11.750 1.5141 0.03266 0.02608 -0.0653 0.0093 1.0000
12.000 1.5165 0.03475 0.02825 -0.0638 0.0090 1.0000
12.250 1.5126 0.03746 0.03109 -0.0621 0.0086 1.0000
12.500 1.5136 0.03984 0.03356 -0.0608 0.0084 1.0000
12.750 1.5187 0.04188 0.03568 -0.0600 0.0083 1.0000
13.000 1.5219 0.04417 0.03804 -0.0591 0.0081 1.0000
13.250 1.5233 0.04671 0.04067 -0.0583 0.0079 1.0000
13.500 1.5237 0.04944 0.04349 -0.0577 0.0077 1.0000
13.750 1.5234 0.05235 0.04649 -0.0571 0.0075 1.0000
14.000 1.5219 0.05545 0.04967 -0.0567 0.0073 1.0000
14.250 1.5198 0.05870 0.05301 -0.0564 0.0072 1.0000
14.500 1.5176 0.06206 0.05645 -0.0563 0.0070 1.0000
14.750 1.5139 0.06567 0.06015 -0.0562 0.0069 1.0000
15.000 1.5091 0.06946 0.06403 -0.0563 0.0067 1.0000
15.250 1.5000 0.07389 0.06855 -0.0565 0.0066 1.0000
15.500 1.4872 0.07894 0.07371 -0.0569 0.0064 1.0000
15.750 1.4716 0.08445 0.07934 -0.0575 0.0063 1.0000
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